The present invention relates to a turbine cooling blade of a gas turbine used for power generation and industry, and more particularly, to cooled turbine blade having an inner hollow structure in which an insert core having an improved structure is accommodated.
A gas turbine used for a power plant is generally arranged as shown in FIG. 11, in which compressed air which is compressed by driving a compressor 2 coaxially provided with a gas turbine 1 is supplied to a combustor 3, fuel is burnt in the liner portion 3a of the combustor 3, and a high temperature combustion gas resulting from the combustion is guided to moving blades 6 through a transition piece 4 and stationary blades 5 of the gas turbine 1, so that the gas turbine 1 delivers work by the rotation of the moving blades 6.
Incidentally, in order to improve the heat efficiency of a gas turbine, it is preferable to increase the turbine inlet temperature and, actually, the turbine inlet temperature is increased for this purpose. As the turbine inlet temperature is increased, it becomes necessary to use a material resistant to high temperature for the combustor 3, the stationary blades 5 and the moving blades 6 of the gas turbine 1 and a heat resistant super-alloy material has been hence been used in gas turbine parts.
Although a heat resistant super-alloy material used in high temperature parts of the turbine has a critical temperature of 800.degree.-900.degree. C. at present, a turbine inlet temperature reaches about 1300.degree. C. which greatly exceeds the critical temperature. Thus, it is essential to employ a cooled blade to which a cooling structure is applied to maintain the reliability of the gas turbine by cooling the blade to the critical temperature.
Air is used as an operating fluid in many cases to cool a blade to its critical temperature, and the air is supplied by being partially extracted from a mid-portion of the compressor 2 or from the path from the outlet of the compressor 2 to the combustor 3. When a larger amount of cooling air is used, the air can further reduce the temperature of the blade. However, the cooling air does not generate an output power until it is collected at a gas passage, and if the cooling air is collected at the gas passage, it reduces the gas temperature. Consequently, the efficiency of a gas turbine is lowered, so that any efficiency improvement due to increased inlet temperature is canceled. Thus, it is an important problem to provide effective cooling using an air flow amount which is as small as possible.
Presently, an air cooled blade as shown in FIG. 12 and FIG. 13 is used to a gas turbine having a turbine inlet temperature of about 1300.degree. C. FIG. 13 is a cross sectional view taken along the line XIII--XIII of FIG. 12. In order to make the following description more understandable, the x-direction, y-direction and z-direction are defined herein as shown in FIGS. 12, 13 and 15.
As shown in FIG. 12 and FIG. 13, an insert core 7 having an inner hollow structure is accommodated in the stationary blade (hereinafter, referred to as a cooling blade main body) 5. The insert core 7 is generally supported in the blade main body 5 at both end portions in a span direction of the blade body 5, and a rib member 5a is normally disposed inside the blade body 5 to support the insert core 7 and further to carry out heat radiation through the rib member 5a. Cooling air 8 is first supplied into the insert core 7, is converted to impingement cooling air 9a by passing through the many impingement holes 9 defined in the insert core 7, and then impinges on the inner surface of the cooling blade main body 5. It is conventionally known that a fluid impinging on a fixed wall at high speed as described above generally has a very high heat transfer coefficient and thus a high cooling efficiency, which is called impingement cooling.
This cooling method is an important cooling technology for managing the cooling effect of the inner surface of the cooling blade main body 5, i.e., an inner surface heat transfer coefficient. Air having once cooled the inner surface of the cooling blade main body 5 then flows out from film holes 10 to the outside in such a fashion that it covers the outer surface of the cooling blade main body 5 in the form of a film. The cooling air film protects the outer surface of the cooling blade main body 5 from high temperature.
Although the cooling air 8 flows to the outside by successively passing through the insert core 7, the impingement holes 9, and the film holes 10 in this order, the cooling air 8 flows from a front edge 12 to a rear edge 13 in a space 11 between the insert core 7 and the cooling blade main body 5. Pin fins 14 are provided in the cooling blade body 5 to increase the heat transfer coefficient and to obtain a fin effect due to the increase of heat transfer area.
FIG. 14 is a view observed from the C direction of FIG. 12 to show a cooling method of a shroud segment 15, in which not only the cooling blade main body 5 but also the shroud segment 15 are subjected to impinge cooling in a gas turbine. That is, many impingement holes 9 are defined to a diaphragm 16, and the shroud segment 15 is cooled by the impingement cooling. Although FIG. 14 shows the impingement cooling applied to a blade root portion, the same cooling technology is also applied to the shroud segment at the extreme end portion of the blade.
Problems of the prior art impingement cooling will be described below. Although impingement cooling is applied to both the cooling blade main body 5 and the shroud segment 15 as described above, since they have the same structure, only the cooling of the cooling blade main body 5 will be described below.
In the gas turbine blade, since the entire surface of the cooling blade main body 5 must be uniformly cooled, the many impingement cooling holes 9 must be defined. As shown in FIG. 15 and FIG. 16, the cooling air 8 flows out from the impingement cooling holes 9 defined in the insert core 7, is converted to impingement cooling air 9a, and impinges on the inner surface of the cooling blade main body 5. It is known that the impingement cooling maximizes the heat transfer coefficient when cooling air impinges perpendicularly on a solid surface to thereby increase the cooling effect. Thus, the states shown in FIG. 15 and FIG. 16 are ideal.
However, in actuality, the cooling air flows as shown in FIG. 17 and FIG. 18. That is, after impinging on the inner surface of the cooling blade main body 5, the impingement cooling air 9a flows from the front edge 12 to the rear edge 13 in the x-direction in the space 11 between the insert core 12 and the cooling blade main body 5 and is converted to cooling air 11a for the space 11. However, this flow of the cooling air 11a interferes with the flow of the cooling air 9a. Therefore, the impingement cooling air 9a does not always impinge perpendicularly on the inner surface of the cooling blade main body 5 and it is impossible to realize the ideal state.
More specifically, when the inner wall of the blade is to be cooled by an impingement jet flow from the insert core 7, as a cross flow flowing between the insert core 7 and the inner wall of the blade is increased with respect to the jet flow, the impingement cooling effect is reduced. Then, in the prior art inside cooling structure, a problem arises in that as the cooling air approaches from the front edge 12 toward the impingement hole row on the downstream side in the rear edge 13, the cross flow of the cooling air which has performed the impingement cooling on the upstream side is increased and it is difficult to obtain a desirable impingement cooling effect.
Further, as the inlet temperature of a gas turbine is increased, an amount of cooling air needed is increases. In particular, when the inlet temperature is 1300.degree. C. or higher, the amount of cooling air is remarkably increased. Moreover, since convection cooling in the inside of the blade is not sufficient, a film cooling method of blowing off cooling air from the film holes 10 defined in the blade surface to the outside of the blade described above must be used together with the convection cooling.
The film cooling method is not only effective for cooling but also prevents a further increase of thermal stress in the metal portion of the blade due to high temperature. As described above, although the application of the film cooling method is effective to cool the cooling blade of a gas turbine, when temperature is further increased, a full coverage film cooling (FCFC) method must be employed to blow off cooling air to the entire surface of the blade.
Since the effect of the film cooling greatly varies depending upon blow-off conditions to a main flow (such as density ratio, mass to quantity of flow ratio, quantity of motion ratio) (that is, there are optimum conditions), there is a possibility that a maximum cooling effect cannot be obtained even by employing the FCFC method. In the case of a turbine stationary blade, there is a large difference in a static pressure on a blade surface onto which cooling air is blown off, depending upon the locations of the blade surface, due to the characteristics of the stationary blade.
Regardless of this fact, the pressure in the space 11 formed between the insert core 7 and the hollow inner wall is kept at a given value in the prior art inside cooling structure. As a result, there is a problem that the pressure of the cooling air, just before it is blown off, cannot be optimized depending upon the location to which the cooling air is blown off and optimum blow-off conditions (such as density ratio, mass to quantity of flow ratio, quantity of motion ratio) cannot be obtained.